Method and means for supporting a ram-jet propellant



` F. L. HAAKE May 16, 1967 METHOD AND MEANS FOR SUPPORTING A RAM-JET PROPELLAN Filed sept. so, 1955 my@ AMNN R.A G m nn 0H. A m MLN w/nK mm @0 n E.

United States Patent 3,319,424 METHOD AND MEANS FUR SUPPRTNG A RAM-.LET PRPL'SLLANT Frederick L. Haal-Ie, Oxnard, Calif., assigner to the United States of America as represented by the Secretary of the Navy Filed Sept. 30, 1965, Ser. No. Ltlzl 3 Claims. (Cl. Gli-255) ABSTRACT F THE DISCLOSURE A device for retaining in position the solid propellant of a ram-jet missile throughout its burning cycle, such device comprising a tubular liner of spring metal enclosing the propellant and having a plurality of inwardlyprojecting tabs which extend toward the front of the projectile, these tabs exerting an inward force on the outer surface of the propellant and wedging into the latter when forces are developed during and subsequent to missile launching.

The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

The present invention relates to a method and means for improving the performance of projectiles of the ramjet type designed to utilize a solid propellant as a source of fuel.

One of the major problems encountered in the design of missiles employing the ram-jet principle of operation is the retention of the solid propellant in place within the combustion chamber until the time that such fuel is completely consumed. It has been found that, in many cases, mechanical separation of the propellant occurs during and subsequent to the launching operation, and almost invariably this results in destruction of the projectile before it has reached its objective. Various expedients have been proposed to bring about a more secure bonding of the propellant to the member on which it is supported, but no one of these solutions has proven to be completely successful.

Among the factors contributing to the difficulties mentioned, apart from the expected vibration of the entire assembly, are (l) the fact that a high inertial force is produced on the propellant due to the rapid acceleration of the projectile immediately subsequent to launching, which force naturally tends to discharge the propellant rearwardly through the nozzle, (2) the fact that the high velocity turbulent ow of gases generated by the cornbustion of the propellant produces a force thereon tending to separate such propellant from the member on which it is carried, and (3) it is not ordinarily feasible to employ radially-extending support members since they interfere with the ilow of air through the combustion chamber and hence tend to degrade missile performance.

In considering solutions to the above problem, it must be borne in mind that no proposal is acceptable unless it permits the projectile to be assembled rapidly and with a minimum of effort. For example, it might be assumed that the solid propellant could be bonded to the interior surface of the missile shell, but this is a diicult assembly operation and precludes the solid propellant from being removed for inspection, as well as limiting access to the other components of the projectile. It would be highly desirable to have available a method and means whereby the solid propellant can be readily and quickly inserted into the missile shell and, once having been so inserted,

remain in place during the launching operation and during that portion of missile flight powered by ram-jet action. It is an objective of the present invention to provide a method and means for accomplishing this objective.

ln a preferred embodiment, the present concept utilizes a thin hollow tube having perforations in the form of a plurality of sets of circumferentially-arranged tabs which are bent inwardly, each set of tabs being spaced longitudinally from the remaining sets. This tube is secured in place within the missile shell to some rigid component of the projectile assembly. When the propellant is inserted in the tube, the tabs are forced or displaced outwardly. As soon as the propellant is ignited, the developed pressure tends to force the propellant in a direction such that the edges of the tabs dig into the material of which the propellant is formed, and this wedging action precludes any further movement of the propellant rearwardly. inasmuch as the tube itself is securely attached to some rigid portion of the missile assembly, the propellant cannot break away to result in a malfunction and consequent possible destruction of the missile.

One object of the present invention, therefore, is to provide a method and means for precluding displacement of the solid propellant forming part of a projectile of the ram-jet type.

A further object of the invention is to provide a method whereby the solid propellant of a ram-jet missile may be readily and quickly inserted in position within the missile shell or casing.

An additional object of the invention is to provide a liner for the solid fuel of a ram-jet missile, such liner consisting of a hollow tube adapted to receive the solid propellant and designed to permit movement of the propellant within the tube only in the direction in which the propellant is originally inserted.

A still further object of the invention is to provide a liner for the solid propellant of a ram-jet missile, such liner being adapted to receive the propellant therewithin and being provided with a plurality of inwardly-projecting tabs designed to contact the outer surface of the propellant and to allow the movement of such propellant past the tabs only in the direction in which the propellant is inserted into the tube, and to preclude relative movement between the tube and propellant in a direction which the latter would normally take when acted upon by forces generated within the missile combustion chamber following ignition of the propellant.

Other objects, advantages and novel features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings wherein:

FIG. l is a partly sectional view, longitudinally broken away, of a projectile of the ram-jet type into which a preferred embodiment of the present invention has been incorporated;

FIG. 2 is an isometric view of one of the components of the invention device, showing the manner in which the latter acts to preclude movement of the propellant in FIG. 3 is an enlarged view of a portion of FIG. l, bringing out more clearly certain details of applicants invention.

Referring now to the drawings, there is shown in FIG. l a ram-jet vehicle of a type now known in the art. This vehicle includes an outer tubular casing, or shell, generally identified by the reference numeral lt?. In order to develop a source of ram air, the forward end of the shell or casing 1) supports an axially-positioned spike 12 which is separated from the interior surface of the shell 10 so as to define an annular air passage therebetween. The

axially-aligned spike 12 may be integrally associated with other portions of the vehicle, such as the guidance system and/or an instrumentation package (not shown) all of which are located in the forward end of the vehicle in a known manner.

At the rear of the projectile shown in FIG. 1 of the drawings is located a nozzle 14 which is associated in the usual fashion with the combustion chamber of the vehicle. The latter contains a body of solid fuel, of any suitable composition, which is designated in FIG. 1 of the drawing by the reference numeral 16. In order to permit the passage of ram air through the vehicle, the solid fuel is pre-formed as best shown in FIG. 2 of the drawings-that is, it is of cruciform cross-section so as to present a plurality of longitudinal air passages from one end of the combustion chamber to the other. It should be understood, however, that the particular form of solid propellant illustrated in the drawings is given merely as an example, and that various other configurations may be employed as long as ram air is capable of being conducted from the forward portion of the vehicle through the combustion chamber to be exhausted through the nozzle 14.

In an arrangement of the type set forth above, considerable difficulty has been encountered in retaining the propellant member 16 in position following the launching operation. This is due to several circumstances, as briefiy mentioned above, one of which being the initial force that is developed by the rapid acceleration of the vehicle immediately subsequent to launching, such force acting to tear the propellant 16 away from its support. Also to be considered is the fact that, as the vehicle travels through the atmosphere, a very intense heat is developed by burning of the propellant, and this tends to weaken the binder which normally forms one of the ingredients of the propellant mixture. As a result, large segments of the propellant tend to break off, and this disintegration of the propellant is liable to block the exhaust nozzle 14 and result in destruction of the projectile. No completely satisfactory solution to this problem has heretofore been proposed.

`In accordance with a preferred embodiment of the present concept, there is provided a liner intermediate the solid propellant member 16 and the inner surface of the tubular shell 10. This liner, best shown in FIG. 2 of the drawings, is identified by the reference numeral 18, and may be composed of some such material as .02 inch thick stainless steel. It is intended to enclose therewithin a major portion of the solid propellant 16, or at least a sufficient amount of the latter so as to preclude any longitudinal movement thereof in a manner now to -be described.

The thin hollow tubular member 18 has formed therein a plurality of perforations, these perforations consisting of a plurality of sets of circumferentially-spaced tabs 20 of generally square configuration, formed by cutting the material of the tubular member 18 on three sides of each square area and then pressing the material of the tab inwardly to result in a condition clearly brought out in FIG. 2 of the drawings. Each set of circumferentially-spaced tabs is spaced apart longitudinally from each other set, and the tabs preferably extend for the entire length of the tube 18.

At the forward end of the tube 18, a member of cruciform configuiration partially closes the open end of the tube. This member, identified in FIGS. 2 and 3 by reference numeral 22, may be made of the same material as that which is utilized to form the tube 18. In the four regions where it lies in contact with the forward edge of the tube 18, it is welded or otherwise securely afiixed to the tube 18- by any conventional process. Member 22 is provided with an axial opening 24 and a further pair of diametrically opposed openings 26 (see FIG. 2) the purpose of which will hereinafter become apparent. It will be noted that the configuration of the cruciform member 22 corresponds to the cross-sectional configuration of the solid propellant member 16, so that, when the propellant 16 is inserted into the tube 18 in a manner to be subsequently decribed, the propellant will be aligned with the cruciform member 22 as clearly brought out in FIG. 2 of the drawings.

As hereinabove stated, the projectile of the drawings contains within its forward portion an instrumentation package (not shown) which is rigidly supported and positioned with respect to the missile body. Extending rearwardly from such instrumentation package is an adaptor identified in FIG. 1 by the reference numeral 28. This adaptor 28 terminates in four radially-extending arms arranged in quadrature, these four radially-extending arms having a common rear surface of planar configuration lying normal to the projectile aXis and substantally coextensive with the cruciform member 22 which partially closes one end of the tubular member 18. This memiber 22 is intended to lie in face-to-face relationship with the planar portion of the adaptor 2S, as again best shown in FIG. 1.

Before proceeding further with a description of the manner in which the various components above mentioned are assembled at the time the missile is made ready for operation, it should be mentioned that the forward portion of the propellant 16 is formualted with a phenolic base, and into this substance is embedded a metallic hub 39 having an axial threaded opening 32 formed therein. This opening 32 is intended to be aligned with the opening 24 in the cruciform end member 22 of FIG. 2. Further, the metallic hub 30 of FIG. l has extending radially therefrom a plurality of arms 34 in the form of a spider- The purpose of these arms 34 is to add to the resistance of the propellant member 16 to longitudinal or axial movement within the shell 10, especially as the propellant burns away and only a small portion thereof remains in the forward end of the tube 18. The hub 30 together with its associated spider is molded into the propellant at the time that the later member is fabricated. The fore portion of the hub 30 is fiush with the end surface of the propellant, so that the latter presents a planar appearance and is suitable for being brought into face-toface relationship with the inner surface of the cruciforrn end closure element 22 of FIG. 2

The propellant 16 is intended to be solely of the endburning type-that is, it burns forwardly only, and no radial burning is desired. Consequently, al1 outer surfaces of the propellant member 16 (except the two end surfaces, and including those surfaces along which air ows in travelling from the forward end of the assembly toward the nozzle 14) are initially coated with some material which inhibits the burning thereof, such material in a preferred embodiment being asbestos of a thickness in the neighborhood of .O3 inch. This so-called inhibitor coating is designated in the drawings by the reference numeral 36. This inhibitor coating 36 appears on those surfaces of the propellant member 16 which are designed to lie adjacent the inner surface of the hollow tubular liner 18, as clearly brought out in FIGS. 1 and 3 of the drawings.

In assembling the elements of the present invention, it is assumed that the nozzle 14 of the projectile and its associated components have been disengaged from the remainder of the projectile body. The components contained within the instrumentation package have also been removed or are not present, permitting access to the interior of the adaptor 28 from the forward portion of the missile. Initially, the liner 18 (to which the cruciform end member 22 has previously been welded) is inserted into the tubular shell 10 from the rear so that the cruciform member 22 lies face-to-face with the lrear planar surface portion of the adaptor 28. The adaptor 28, as shown in FIG. 1, has three openings formed in the rear planar surface portion thereof, these three openings being aligned with the openings 24 and 26 of the cruciform member 22. When the liner 13 has been inserted and is in position as above described, a pair of screws 37 are inserted into the openings 26 from the interior of the liner 18 and are forced into their associated openings in the adaptor 2S. The inner surface of the cruciform member 22 surrounding the openings 26 is countersunk as shown in FIG. l to receive the slant heads of the screws 37. When these screws which may be of the self tapping variety have been hammered into place, the cruciform member 22, and hence the liner 18 is securely held in position against the rear surface portion 3S of the adaptor Z8.

The propellant member 16 is now inserted into the liner 18, with the end thereof containing the hub 30 and oriented forwardly. It will be noted that as the propellant member 16 enters the tubular 18, the forward edge of propellant depresses each one of the tabs 20 in succession. In other words, each tab 20 is forced outwardly and bent back toward the original position it held before it was depressed. However, since the material of which the liner 1S is constituted possesses a certain amount of resiliency, there is a spring action involved which causes each one of the tabs 20 to press strongly upon the inhibitor coating 36 which overlies the outer surface of each cruciform portion of propellant member 16. Referring to FIG. 3, it will be seen that each one of the tabs 2@ presses downwardly upon this inhibitor coating.

When entry of the propellant 16 has been completed, the axial aperture 32 of the hub 30 will be .aligned with the opening 24 in the cruciform member 22 and also aligned with the axial opening in the planar surface 38 in the adaptor 28. A bolt 40 is then inserted through this axial opening in the surface portion 38 of the adaptor 23 `and through the opening 24 of cruciform member 22 into the opening 32 of the hub 30. When this bolt 40 has been tightened, it secures the liner 18 from any longitudinal motion with respect to the adaptor 28 and hence with respect to the shell 10,

Reference to FIG. 3 of the drawings will show that the sharp edges of each one of the tabs 20 digs into the asbestos (or inhibitor) coating 36 on the propellant member 16. Consequently, any force received by the propellant member 16 which tends to displace it rearwardly of the missile `shell will cause each one of the tabs to dig further into this inhibitor coating 36, and the wedging action of all of the tabs 20, when taken together, completely precludes any longitudinai motion of the propellant member 16 with respect to the liner 18. Since the liner is securely held to the adaptor 28 by means `of the bolt 40, no displacement of either the liner or propellant can occur even though a considerable portion of the propellant is consumed at the time such a separation wo-uld otherwise tend to take place.

Inasmuch as the propellant member 16 is of the endburning type, it becomes progressively shorter during flight of the projectile. Since there is no necessity for the liner 18 to extend beyond that end of the propellant member 16 lat which burning occurs, the composition of the liner 18 may be chosen to burn away along with the propellant 16. It has been found that a stainless steel member of approximately the thickness given above will be consumed in this fashion due to the intense heat which is produced by burning of the propellant during projectile flight.

The hollow tubular liner 18 has been shown in FIGS. l and 3 of the drawings as being spaced apart from the surface of the missile shell 10 inner. Under certain conditions, and depending upon the thickness and resiliency of the material making up the liner, insertion of the propellant member 16 into the liner in the manner above described will cause the latter to be somewhat deformed from a perfectly annular conguiration so that it will have a somewhat increased diameter in the regions where the tabs appear. This might cause the liner 18 to directly contact the inner surface of the shell 10 at such points, but this will have no adverse effect on the manner in which the device operates.

The length of the liner 18, Vand the number of tabs, which are formed therein, will depend to a large extent upon the amount of retaining action which is either necessary Ior desirable to preclude separation of the propellant during launching and subsequent missile flight. 'Such dimensions and characteristics are readily determined experimentally, -although obviously an excess number of tabs, or greater than necessary length of the liner, will not detract from missile performance.

Obviously many modications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically described.

I claim:

1. In an elongated projectile of the ram-jet type powered by a solid propellant 4all or a portion of which has a tendency to be longitudinally displaced in position with respect to the projectile shell as a result of forces developed during, and subsequent to, the launching operation, which displacement, when allowed to occur, resulting in an irregularity in t-he burning cycle of said propellant .and a possible destruction of the projectile before it has reached its objective, the improvement which comprises:

means for retaining said propellant in position throughout its burning cycle, said means including:

a liner of generally tubular configuration composed of spring metal and disposed intermediate said solid propellant ,and the shell of said projectile, said liner being secured to the projectile body and enclosing at least a major portion of said solid propellant therewithin;

said liner having a plurality of integrally-formed inwardly-extending projections exerting by spring action an inward force on the outer surface of said solid propellant;

said inwardly-extending projections being in the form of a plurality of sets of circumferentiallyspaced tabs, each set of tabs being longitudinally separated from each other throughout the length of said liner;

each tab of each of said sets extending both inwardly of said liner and toward the front of said projectile, so as to create a wedging action between such tab and the outer surface of said solid propellant thereby precluding `any appreciable displacement of the latter when the said forces are developed during and subsequent t0 the launching operation.

2. In the assembly of a solid-propellant ram-jet missile, such missile being designed with a liner for retaining such propellant in position regardless of forces tending to dislodge the propellant both during and subsequent to the launching operation, such liner being in the form of a tubular member of resilient material having formed therein -a plurality of extensions projecting angularly with components both inwardly and axially of said tubular member, ya method which comprises the steps of:

inserting said liner into the shell of said missile so that vsaid extensions project both inwardly and toward the front of said missile;

securing the lforward end of said liner to the body of said missile to preclude any relative movement therebetween;

inserting said solid propellant into said liner from the rear of the latter so that, as the propellant passes in turn each of said extensions, the latter are forced outwardly but continue to exert an inward pressure on said propellant;

whereby any subsequent force on said propellant exerted rearwardly of said missile will cause the outer edge of each of said extensions to Wedge itself into the material of which said propellant is composed and prevent any Vappreciable relative movement between said propellant and said liner and hence between said propellant and the body of said missile.

3. The method 0f claim 2, in which the circumferential surface of said propellant is provided with an inhibitor coating, said extensions becoming wedged into such inhibitor coating upon the exertion `of a force on said pro- 10 pellant rearwardly of said missile following the insertion of said propellant into said liner.

References Cited by the Examiner UNITED STATES PATENTS 2/1956 ll/l957 l/l958 5/1963 10/1963 l/l964 2/1965 Dickinson 60-255 Topinka 60-255 Tarr 60-255 Brewer 60--255 De Fries et al 60255 Damon et al. Y

Adelman 60--255 CARLTON R. CROYLE, Primary Examiner. 

1. IN AN ELONGATED PROJECTILE OF THE RAM-JET TYPE POWERED BY A SOLID PROPELLANT ALL OR A PORTION OF WHICH HAS A TENDENCY TO BE LONGITUDINALLY DISPLACED IN POSITION WITH RESPECT TO THE PROJECTILE SHELL AS A RESULT OF FORCES DEVELOPED DURING, AND SUBSEQUENT TO, THE LAUNCHING OPERATION, WHICH DISPLACEMENT, WHEN ALLOWED TO OCCUR, RESULTING IN AN IRREGULARITY IN THE BURNING CYCLE OF SAID PROPELLANT AND A POSSIBLE DESTRUCTION OF THE PROJECTILE BEFORE IT HAS REACHED ITS OBJECTIVE, THE IMPROVEMENT WHICH COMPRISES: MEANS FOR RETAINING SAID PROPELLANT IN POSITION THROUGHOUT ITS BURNING CYCLE, SAID MEANS INCLUDING: A LINER OF GENERALLY TUBULAR CONFIGURATION COMPOSED OF SPRING METAL AND DISPOSED INTERMEDIATE SAID SOLID PROPELLANT AND THE SHELL OF SAID PROJECTILE, SAID LINER BEING SECURED TO THE PROJECTILE BODY AND ENCLOSING AT LEAST A MAJOR PORTION OF SAID SOLID PROPELLANT THEREWITHIN; SAID LINER HAVING A PLURALITY OF INTEGRALLY-FORMED INWARDLY-EXTENDING PROJECTIONS EXERTING BY SPRING ACTION AN INWARD FORCE ON THE OUTER SURFACE OF SAID SOLID PROPELLANT; SAID INWARDLY-EXTENDING PROJECTIONS BEING IN THE FORM OF A PLURALITY OF SETS OF CIRCUMFERENTIALLYSPACED TABS, EACH SET OF TABS BEING LONGITUDINALLY SEPARATED FROM EACH OTHER THROUGHOUT THE LENGTH OF SAID LINER; EACH TAB OF EACH OF SAID SETS EXTENDING BOTH INWARDLY OF SAID LINER AND TOWARD THE FRONT OF SAID PROJECTILE, SO AS TO CREATE A WEDGING ACTION BETWEEN SUCH TAB AND THE OUTER SURFACE OF SAID SOLID PROPELLANT THEREBY PRECLUDING ANY APPRECIABLE DISPLACEMENT OF THE LATTER WHEN THE SAID FORCES ARE DEVELOPED DURING AND SUBSEQUENT TO THE LAUNCHING OPERATION. 